Ceramic matrix composite component, turbine system and fabrication process

ABSTRACT

A ceramic matrix composite component, turbine system and fabrication process are disclosed. The ceramic matrix composite (CMC) component includes a CMC material, an environmental barrier coating (EBC) on the CMC material, and a hard wear coating applied over the EBC. The turbine system includes a rotatable CMC component having a hard wear coating, and a stationary turbine component, the stationary turbine component having an abradable coating arranged and disposed to be cut by the silicon carbide material. The fabrication process includes positioning the rotatable CMC a pre-determined distance from the stationary turbine component and rotating the rotatable CMC component. The hard wear coating on the rotatable CMC component cuts the abradable coating on the stationary turbine component.

STATEMENT REGARDING FEDERALLY SPONSORED RESEARCH OR DEVELOPMENT

This invention was made with government support under contract numberDE-FC26-05NT42643 awarded by the Department of Energy. The governmenthas certain rights in the invention.

FIELD OF THE INVENTION

The present invention is directed to manufactured components and aprocess of using manufactured components. More particularly, the presentinvention is directed to a wear coating applied to ceramic matrixcomposite (CMC) components.

BACKGROUND OF THE INVENTION

Gas turbines are continuously being modified to provide increasedefficiency and performance. These modifications include the ability tooperate at higher temperatures and under harsher conditions, which oftenrequires material modifications and/or coatings to protect componentsfrom such temperatures and conditions. As more modifications areintroduced, additional challenges are realized.

One modification to increase performance and efficiency involvesminimizing a gap between a turbine bucket tip and a turbine shroud. Theminimized gaps lead to rubs between the bucket tip and the shroud duringcertain power transients. An environmental barrier coating (EBC) used onCMC gas turbine bucket tips can be damaged by these rubs.

Damage to the EBC on a turbine bucket may expose an underlying CMC tohigh temperature combustion gasses leading to increased volatilization.Increasing the EBC thickness on the bucket increases the weight of thebucket and decreases the efficiency and performance of the turbine.Furthermore, opening the gap to minimize rubs also decreases efficiencyand performance.

A manufactured component and process of using manufactured componentsnot suffering from the above drawbacks would be desirable in the art.

BRIEF DESCRIPTION OF THE INVENTION

In an exemplary embodiment, a ceramic matrix composite (CMC) componentincludes a CMC material, an environmental barrier coating applied overthe CMC material, and a hard wear coating applied over the environmentalbarrier coating.

In another exemplary embodiment, a turbine system includes a rotatableCMC component having a CMC material, an environmental barrier coating(EBC) applied over the CMC material, and a hard wear coating appliedover the EBC. The turbine system further includes a stationary turbinecomponent having an abradable coating arranged and disposed adjacent tothe rotatable CMC component, to be cut by the hard wear coating of therotatable turbine component.

In another exemplary embodiment, a fabrication process includespositioning a rotatable CMC component a pre-determined distance from astationary turbine component, the stationary turbine componentcomprising an abradable material. The fabrication process furtherincludes rotating the rotatable CMC component, wherein, a hard wearcoating on the rotatable CMC component cuts the abradable material ofthe stationary turbine component.

Other features and advantages of the present invention will be apparentfrom the following more detailed description of the preferredembodiment, taken in conjunction with the accompanying drawings whichillustrate, by way of example, the principles of the invention.

BRIEF DESCRIPTION OF THE DRAWINGS

FIG. 1 is a perspective view of a turbine bucket.

FIG. 2 is a schematic view of a turbine system.

FIG. 3 is a schematic view of a rotating part to stationary partinterference pattern.

Wherever possible, the same reference numbers will be used throughoutthe drawings to represent the same parts.

DETAILED DESCRIPTION OF THE INVENTION

Provided is an exemplary ceramic matrix composite component, turbinesystem and fabrication process. Embodiments of the present disclosure,in comparison to methods and products not utilizing one or more featuresdisclosed herein, utilize decreased environmental barrier coating (EBC)material, decrease damage to the EBC, decrease repairs, decrease ceramicmatrix composite (CMC) volatilization, decrease tip clearance, increaseturbine efficiency, or a combination thereof.

Referring to FIG. 1, in one embodiment, a rotatable CMC component 101includes a CMC material 102, an EBC 104 applied over the CMC material102, and a hard wear coating 106 applied over the EBC 104. The rotatableCMC component 101 is any suitable component that may experiencevolatilization and/or rub wear such as, but not limited to, a blade orbucket. The bucket may be either an unshrouded bucket or a shroudedbucket. The hard wear coating 106 is a material such as, but not limitedto silicon carbide (SiC), SiO₂, cubic boron nitride (CBN), orcombinations thereof. Examples of CMC material 102 include, but are notlimited to, carbon-fiber-reinforced carbon (C/C),carbon-fiber-reinforced silicon carbide (C/SiC),silicon-carbide-fiber-reinforced silicon carbide (SiC/SiC), andalumina-fiber-reinforced alumina (Al₂O₃/Al₂O₃). The CMC's have increasedelongation, fracture toughness, thermal shock, dynamical loadcapability, and anisotropic properties as compared to a monolithicceramic structure. However, without an environmental coating, the CMC'smay volatilize during gas turbine operation.

For example, at temperatures above 1500° F. water vapor chemicallyreacts with the CMC material 102. The water vapor reacts with siliconand carbon in the CMC material 102 to produce silicon hydroxide (SiOH)and carbon dioxide (CO₂), respectively. The SiOH and CO₂ formed by thereaction between the water vapor and the CMC material 102 slowly gassesoff, or volatilizes. Over many hours of operation above 1500° F. the CMCmaterial 102 is removed from the outside in.

The EBC 104 protects the CMC material 102 from water vapor, heat, andother combustion gases which may cause the volatilization ordeterioration of the CMC material 102. In one embodiment, the EBC 104reduces or eliminates the occurrence of the chemical reaction betweenthe water vapor and the CMC material 102. The EBC 104 may be anysuitable material for protecting the CMC material 102 from the hot gasesof combustion. Suitable EBC materials include, but are not limited to,barium strontium alumino silicate (BSAS), mullite, yttria-stabilizedzirconia, and combinations thereof.

Referring to FIG. 1, in one embodiment, the hard wear coating 106 formsan outermost layer of the rotatable CMC component 101, being appliedover one or more EBCs 104. The hard wear coating 106 is applied over anysuitable portion of the rotatable CMC component 101 that experienceswear during a rub event. Suitable portions of the rotatable CMCcomponent 101 include, but are not limited to, a tip portion 103, aplatform interface, the contact point between an airfoil and a shroud,the contact point between the airfoil and shroud rails, or a combinationthereof. Additionally, the hard wear coating 106 is applied with anysuitable thickness to reduce or eliminate a loss of the EBC 104 duringdeep rubs. Suitable thicknesses include, but are not limited to, betweenabout 0.1 mils and about 4 mils, between about 0.5 and about 3 mils,between about 1 mil and about 2.5 mils, or any combination,sub-combination, range, or sub-range thereof. For example, in oneembodiment, the hard wear coating 106 extends across the entire widthand length of the tip portion 103. In another example, the hard wearcoating 106 is applied around an edge of the tip portion 103 and down toa radially extending surface of the rotatable CMC component 101 toprovide increased protection from severe conditions.

The hard wear coating 106 is applied over the EBC 104 through anysuitable deposition process. For example, one suitable depositionprocess is physical vapor deposition (PVD). PVD condenses a vaporizedform of the hard wear coating 106 on the EBC 104 of the rotatable CMCcomponent 101 to form a hard thin hard wear coating 106. Other suitabledeposition process include, but are not limited to, chemical vapordeposition (CVD), air plasma spray (APS), combustion spraying withpowder or rod, slurry coating, sol gel, electrophoretic deposition, tapecasting, or a combination thereof. In one embodiment, SiC is used as thehard wear coating 106. The SiC adheres well because a coefficient ofthermal expansion (CTE) of the SiC is well matched to the CTE for boththe CMC material 102 and the EBC 104. The EBC 104 may be applied to andin contact with the CMC material 102. The hard wear coating 106 may beapplied to and contact the EBC 104.

Referring to FIG. 2, in one embodiment, a turbine system 100 includesthe rotatable CMC component 101 such as a turbine bucket attached to arotatable member 107 such as a turbine disc. The CMC component 101 andthe rotatable member 107 extend away from a central point 131 such as arotatable shaft. In a further embodiment, a plurality of rotatable CMCcomponents 101 are positioned around the rotatable member 107, andextend away from the central point 131. In one embodiment, a stationaryturbine component 120 forms a perimeter around the rotatable member 107,the CMC component 101 being between the rotatable member 107 and thestationary turbine component 120. The stationary turbine component 120is centered around the central portion 131, sharing a common centralpoint with the rotatable member 107. The tip portion 103 of therotatable CMC component 101 ideally forms a seal with the stationaryturbine component 120.

An abradable coating 122 on the stationary turbine component 120 isarranged and disposed to be cut by the rotatable CMC component 101. Theabradable coating 122 is any suitable coating based upon stationaryturbine component 120 material, and operating temperature. Suitableabradable coatings 122 for a metallic stationary turbine compound 120 atlower temperatures (up to approximately 2200° F.) include, but are notlimited to, metal abradables having the general form MCrAlY, such asNiCrAlY, CoCrAlY, FeCrAlY, or combinations thereof. In one embodiment,the metal abradable is applied by air plasma spray, wire arc andcombustion spraying, high velocity oxy-fuel (HVOF), or combinationsthereof.

Suitable abradable coatings 122 for the metallic stationary turbinecomponent 120 at higher temperatures (at least approximately 2200° F.)include, but are not limited to, ceramic abradables such as partiallystabilized zirconia, with yttria as the stabilizer (7 or 8 YSZ). In oneembodiment, the metal abradables for lower temperature applicationsand/or the ceramic abradables for higher temperature applications,initially contain a polyester component during application. Thepolyester component is burned out by heating in air to leave behindporosity. The porosity makes the abradable coating 122 compliant with arotating CMC component 101 during a rub. In one embodiment, at least oneof the metal abradable coating and the ceramic abradable coating furtherincludes boron nitride as a solid lubricant.

Suitable abradable coatings 122 for a ceramic stationary turbinecomponent 120 include, but are not limited to, silicates. For example,in one embodiment, the suitable abradable coating 122 for the ceramicstationary turbine component 120 operating at lower temperatures (up toapproximately 2200° F.) is barium strontium alumino silicate (BSAS). Asanother example, in one embodiment, the suitable abradable coatings 122for the ceramic stationary turbine component 120 operating at highertemperatures (at least approximately 2201° F.) include high rare earthcontent silicates, such as Yb₂O₃Si₂O₇.

In one embodiment, the silicate abradable coatings 122 include apolyester component during application. The polyester component isburned out by heating in air to leave behind porosity which contributesto compliance. In one embodiment, the silicate abradable coating 122 isapplied by air plasma spray, slurry coating, sol gel, electrophoreticdeposition, tape casting, or a combination thereof.

Referring to FIG. 3, in one embodiment, the fabrication process 200includes positioning the rotatable member 107 a pre-determined distance105 from the stationary turbine component 120. The rotatable member 107is rotated around the central point 131, rotating the CMC component 101.During power transients the pre-determined distance 105 fluctuates,resulting in contact between the tip portion 103 and the abradablecoating 122 of the stationary turbine component 120. The hard wearcoating 106 cuts into the abradable coating 122, forming a pathway 201to reduce contact during further power transients. In one embodiment,the direction of rotation 203 is clockwise. In another embodiment, thedirection of rotation 203 is counter-clockwise (not shown).

Referring to FIG. 1, in one embodiment, a hardness value of the hardwear coating 106 is higher than the hardness value of the EBC 104. Thehardness value represents an ability of a material to withstand contactwithout being damaged. For example, the hard wear coating 106 having thehigher hardness value than the EBC 104 is able to withstand deeper rubsthan the EBC 104.

In one embodiment, the hard wear coating 106 having the higher hardnessvalue is applied over the EBC 104 to protect the EBC 104 from beingdamaged when energy transients create contact between the rotating CMCcomponent 101 and the stationary turbine component 120. In oneembodiment, the hard wear coating 106 is SiC, the SiC having a knoopvalue (hardness) of 2480 and the EBC 104 having a knoop value of lessthan 2480. In one embodiment, the EBC 104 has a knoop value of betweenapproximately 500 to 1800, depending on the architecture of the EBC 104.Architectural elements include the number of layers, the thickness ofeach layer, layer composition, the application process and the heattreat used.

For example, in one embodiment, the EBC 104 is only able to survivecontact between the tip portion 103 and the abradable coating 122 of upto about 10 mils. Gas turbine rotating parts incur contact againststatic parts well in excess of 10 mils. Contact above 10 milsdeteriorates the EBC 104 structure and function while possiblysubjecting the base material to environmental attack. To protect the EBC104, the hard wear coating 106 is applied over the EBC 104, decreasingor eliminating damage to the EBC 104 on deep rubs.

Typical turbine bucket running conditions include water vapor, hot gasesof combustion at temperatures in the range of approximately 2,000° F. toapproximately 3,000° F. The hard wear coating 106 gradually volatizesunder these conditions, exposing the EBC 104. Applying a minimalthickness of the hard wear coating 106 minimizes a gap between the EBC104 on the rotatable CMC component 101 and the abradable coating 122 ofthe stationary turbine component 120. Decreasing the gap increases theefficiency of a system by preventing increased air flow between therotatable CMC component 101 and the stationary turbine component 120.

The time it takes for the sacrificial hard wear coating 106 to volatizeunder operating conditions is a break-in period. In one embodiment, thebreak-in period is up to about 100 hours. In one embodiment, thebreak-in period is at least about 100 hours. The break-in period duringwhich the hard wear coating 106 remains on the EBC 104 in the tipportion 103 is sufficient so that the hard wear coating 106 is presentto abrade during power transients resulting in deep rubs.

In one embodiment, the break-in period includes a constantpre-determined speed of rotation for a predetermined time. In oneembodiment, the break-in period includes a pre-determined fluctuationand/or series of fluctuations in speed of rotation. The pre-determinedspeed of rotation or fluctuations in speed of rotation are set to inducepower transients during the break-in period. During the break-in period,the power transients of the fabrication process 200 cause the hard wearcoating 106 to contact, and cut, the abradable material 122. The cut bythe hard wear coating 106 forms a pathway 201 in the abradable material122.

The hard wear coating 106 gradually volatilizes throughout the break-inperiod, creating the pathway 201 between the stationary abradablematerial and the tip portion 103. The break-in period minimizes thepathway 201 between the stationary component and the tip portion 103around which gas may leak, without damaging the EBC 104. Gas leakage isminimized due to the cutting character of the hard wear coating 106 andits minimal thickness. Forming the minimized space between the pathway201 and the tip portion 103, increases the efficiency of the turbinesystem 100. Additionally, the pathways 201 formed in the stationaryabradable material by the cuts from the hard wear coating 106 decreasedamage to the EBC 104 due to rubs after the hard wear coating 106 hasvolatilized. The undamaged EBC 104 provides the CMC material 102 withincreased volatilization protection.

While the invention has been described with reference to a preferredembodiment, it will be understood by those skilled in the art thatvarious changes may be made and equivalents may be substituted forelements thereof without departing from the scope of the invention. Inaddition, many modifications may be made to adapt a particular situationor material to the teachings of the invention without departing from theessential scope thereof. Therefore, it is intended that the inventionnot be limited to the particular embodiment disclosed as the best modecontemplated for carrying out this invention, but that the inventionwill include all embodiments falling within the scope of the appendedclaims.

What is claimed is:
 1. A ceramic matrix composite component, comprising:a ceramic matrix composite material; an environmental barrier coatingapplied over the ceramic matrix composite material; and a hard wearcoating applied over the environmental barrier coating.
 2. The ceramicmatrix composite component of claim 1, wherein the hard wear coatingfurther comprises silicon carbide.
 3. The ceramic matrix compositecomponent of claim 1, wherein the hard wear coating is the outermostlayer.
 4. The ceramic matrix composite component of claim 1, wherein thehard wear coating coats an outer surface of the component.
 5. Theceramic matrix composite component of claim 1, wherein the hard wearcoating is applied by physical vapor deposition.
 6. The ceramic matrixcomposite component of claim 1, wherein the hard wear coating is appliedby chemical vapor deposition.
 7. A turbine system, comprising: arotatable ceramic matrix composite bucket attached to a disc, therotatable ceramic matrix composite bucket having a ceramic matrixcomposite material, an environmental barrier coating applied over theceramic matrix composite material, and a hard wear coating applied overthe environmental barrier coating; and a stationary turbine shroud, thestationary turbine shroud having an abradable coating arranged anddisposed to be cut by the hard wear coating of the rotatable ceramicmatrix composite bucket.
 8. The turbine system of claim 7, wherein thestationary turbine shroud forms a perimeter around the rotatable bucket.9. The turbine system of claim 7, wherein a plurality of rotatableceramic matrix composite buckets are attached circumferentially to theturbine disc.
 10. A fabrication process, comprising: positioning arotatable ceramic matrix composite component a pre-determined distancefrom a stationary turbine component, the stationary turbine componenthaving an abradable coating and the rotatable component having a hardwear layer applied over an environmental layer; and rotating therotatable ceramic matrix composite component; wherein, the hard wearlayer on the rotatable ceramic matrix composite component cuts theabradable coating on the stationary turbine component.
 11. Thefabrication process of claim 10, wherein power transients cause therotatable ceramic matrix composite component to contact the abradablecoating.
 12. The fabrication process of claim 10, wherein the cut by therotatable ceramic matrix composite component forms a pathway in theabradable coating.
 13. The fabrication process of claim 10, wherein theprocess further comprises a break-in period.
 14. The fabrication processof claim 13, wherein the break-in period is up to about 100 hours. 15.The fabrication process of claim 13, wherein the break-in period is atleast 100 hours.
 16. The fabrication process of claim 10, wherein thehard wear layer volatilizes after predetermined exposure to elevatedtemperature.
 17. The fabrication process of claim 16, wherein theelevated temperature comprises at least 1,500° F.
 18. The fabricationprocess of claim 16, wherein the elevated temperature is at least 2,900°F.
 19. The fabrication process of claim 16, wherein the hard wearcoating volatilization exposes the environmental barrier coating. 20.The fabrication process of claim 10, wherein the environmental barriercoating has a hardness value of less than that of the hard wear coating.